Airfoils for gas turbine engines

ABSTRACT

An airfoil for a turbine engine includes a body defining a pressure side and a suction side, the body extending between a leading edge and a trailing edge and also extending radially to define an outer edge. A radially-extending plenum can be disposed at the outer edge of the airfoil. A flared portion can be disposed at the outer edge of the body extending from the suction side.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a continuation of U.S. patent application Ser. No.17/148,156, filed Jan. 13, 2021, now allowed, which is incorporatedherein by reference in its entirety.

FEDERALLY SPONSORED RESEARCH

This invention was made with Government support under W58RGZ-16-C-0047awarded by the U.S. Army. The Government has certain rights in thisinvention.

FIELD

The present subject matter relates generally to airfoils for gas turbineengines.

BACKGROUND

Gas turbine engines include a compressor that provides pressurized airto a combustor wherein the air is mixed with fuel and ignited forgenerating hot combustion gases. These gases flow downstream to one ormore turbines that extract energy therefrom to power the compressor andprovide useful work such as powering an aircraft in flight. In aturbofan engine, which typically includes a fan placed at the front ofthe core engine, a high pressure turbine powers the compressor of thecore engine. A low pressure turbine is disposed downstream from the highpressure turbine for powering the fan. Each turbine stage commonlyincludes a stationary turbine nozzle followed by a turbine rotor.

Each turbine rotor carries a circumferential array of airfoil-shapedturbine blades adapted to extract energy from the combustion gasesexiting the core. These blades are typically constructed by casting fromhigh-temperature resistant alloys (e.g. “superalloys”). The first rotorstage, immediately downstream of the combustor, is usually internallycooled and has a hollow interior with one or more serpentine passages,film cooling holes, trailing edge slots or holes, and the like. Thesubsequent rotor stages are not subject to the extreme high temperatureof the first stage and thus may not require cooling or the same degreeof cooling. To reduce the weight of the later-stage airfoils, they ofteninclude a hollowed-out portion referred to as a plenum.

The interface between the blade and shroud of the gas turbine engineaffects the performance of the gas turbine engine. Accordingly, there isa need for a blade having low weight while maintaining high strength andhaving improved aerodynamic performance.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appended Figs.,in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine incorporating an exemplary embodiment of a turbine sectionaccording to an exemplary embodiment of the present disclosure.

FIG. 2 is perspective view of a portion of an airfoil for a gas turbineengine according to an exemplary embodiment of the present disclosure.

FIG. 3 is a cross-sectional view of the airfoil as seen along Line A-Ain FIG. 2 in accordance with an exemplary embodiment of the presentdisclosure.

FIG. 4 is a cross-sectional view of a portion of an airfoil disposedwithin a gas turbine engine according to an exemplary embodiment of thepresent disclosure.

FIG. 5 is a cross-sectional view of a portion of an airfoil disposedwithin a gas turbine engine according to an exemplary embodiment of thepresent disclosure.

FIG. 6 is an enlarged view of a portion of the airfoil and gas turbineengine as seen in Circle C in FIG. 5 in accordance with an exemplaryembodiment of the present disclosure.

FIG. 7 is an enlarged view of a portion of the airfoil and gas turbineengine as seen in Circle B in FIG. 4 in accordance with an exemplaryembodiment of the present disclosure.

FIG. 8 is a top view of an airfoil as seen from an outer edge accordingto an exemplary embodiment of the present disclosure.

FIG. 9 is a cross-sectional view of the airfoil of FIG. 8 as seen alongLine B-B according to an exemplary embodiment of the present disclosure.

FIG. 10 is a top view of an airfoil as seen from an outer edge accordingto an exemplary embodiment of the present disclosure.

FIG. 11 is a cross-sectional view of the airfoil of FIG. 10 as seenalong Line C-C according to an exemplary embodiment of the presentdisclosure.

FIG. 12 is a top view of an airfoil as seen from an outer edge accordingto an exemplary embodiment of the present disclosure.

FIG. 13 is a cross-sectional view of the airfoil of FIG. 12 as seenalong Line D-D according to an exemplary embodiment of the presentdisclosure.

FIG. 14 is a top view of an airfoil as seen from an outer edge accordingto an exemplary embodiment of the present disclosure.

FIG. 15 is a cross-sectional view of the airfoil of FIG. 14 as seenalong Line E-E according to an exemplary embodiment of the presentdisclosure.

FIG. 16 is a top view of an airfoil as seen from an outer edge accordingto an exemplary embodiment of the present disclosure.

FIG. 17 is a cross-sectional view of the airfoil of FIG. 16 as seenalong Line F-F according to an exemplary embodiment of the presentdisclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, affixing, or attaching, as well as indirect coupling,affixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

In general, rotor blades (referred to herein as airfoils) are used ingas turbine engines to convert between axial and rotational forces.Aerodynamic characteristics of airfoils can affect engine performanceand efficiency. Airfoils described in accordance with embodiments hereincan generally include flared portions along the suction surfacesidewall. The flared portion can extend from the outer edge of theairfoil in an axial direction of the gas turbine engine, creating adesirable aerodynamic effect which increases engine performance andefficiency. In some embodiments, a plenum disposed at the outer edge ofthe airfoil can have a multi-segmented sidewall to further increasedesirable aerodynamic characteristics while reducing unnecessary weightof the airfoil. In addition, one or more cooling elements, such asrecesses and/or cooling holes, can extend from the radially outersurface of the airfoil to a cooling cavity in order to further enhanceaerodynamic performance of the gas turbine engine.

Inclusion of an airfoil in accordance with such configuration including,e.g., a flared portion along the suction surface sidewall may result ina more aerodynamically capable airfoil, permitting increased engineperformance and efficiency. Further, inclusion of an airfoil inaccordance with other aspects of such a configuration may result in anairfoil having a reduced weight and increased cooling characteristics.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1 , the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference), a radial direction R, and a circumferential direction (i.e.,a direction extending about the axial direction A; not depicted). Ingeneral, the turbofan 10 includes a fan section 14 and a core turbineengine 16 disposed downstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. The compressorsection, combustion section 26, and turbine section together define acore air flowpath 37 extending from the annular inlet 20 through the LPcompressor 22, HP compressor 24, combustion section 26, HP turbinesection 28, LP turbine section 30 and jet nozzle exhaust section 32. Ahigh pressure (HP) shaft or spool 34 drivingly connects the HP turbine28 to the HP compressor 24. A low pressure (LP) shaft or spool 36drivingly connects the LP turbine 30 to the LP compressor 22.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades40, e.g., in unison. The fan blades 40, disk 42, and actuation member 44are together rotatable about the longitudinal axis 12 by LP shaft 36across a power gear box 46. The power gear box 46 includes a pluralityof gears for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1 , the disk 42 iscovered by rotatable front spinner cone 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated that forthe embodiment depicted, the nacelle 50 is supported relative to thecore turbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50extends over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the LPcompressor 22. The ratio between the first portion of air 62 and thesecond portion of air 64 is commonly known as a bypass ratio. Thepressure of the second portion of air 64 is then increased as it isrouted through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to an inner casing (not shown) and HP turbine rotor blades 70that are coupled to the HP shaft or spool 34, thus causing the HP shaftor spool 34 to rotate, thereby supporting operation of the HP compressor24. The combustion gases 66 are then routed through the LP turbine 30where a second portion of thermal and kinetic energy is extracted fromthe combustion gases 66 via sequential stages of a plurality of LPturbine rotor blades 72. The plurality of LP turbine rotor blades 72drive the LP shaft or spool 36, thus causing the LP shaft or spool 36 torotate. Such thereby supports operation of the LP compressor 22 and/orrotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration. For example, in other exemplary embodiments, theturbine fan engine 10 may instead be configured as any other suitableturbomachine including, e.g., any other suitable number of shafts orspools, and excluding, e.g., the power gearbox 46 and/or fan 38, etc.Accordingly, it will be appreciated that in other exemplary embodiments,the turbofan engine 10 may instead be configured as, e.g., a turbojetengine, a turboshaft engine, a turboprop engine, etc., and further maybe configured as an aeroderivative gas turbine engine or industrial gasturbine engine.

FIG. 2 is a perspective view of a radially outer portion of an exemplaryairfoil 200 in accordance with an exemplary embodiment described herein.Airfoils described herein can be one of a plurality of rotor blades of aturbine on a gas turbine engine, such as the turbofan jet engine 10previously described with respect to FIG. 1 . Alternatively, the airfoilcan be one of a plurality of rotor blades of a compressor on a gasturbine engine. In yet other embodiments, the airfoil may be associatedwith another piece of equipment on an engine. In the installed state,the airfoil can extend from a central hub of the turbine or compressorradially outward toward a shroud of the gas turbine engine.

The airfoil 200 depicted in FIG. 2 , generally includes a body 202having a suction surface 204 (sometimes referred to as the uppersurface) and a pressure surface 206 (sometimes referred to as the lowersurface). The suction surface 204 and pressure surface 206 can generallycorrespond with the shape and operational movement, e.g., rotationaldirection, of the airfoil 200, with the suction surface 204 beinggenerally associated with higher velocity and lower static pressurewhile the pressure surface 206 is associated with comparatively higherstatic pressure. The suction surface 204 and pressure surface 206 candefine opposite sides of the airfoil 200 and can extend between an outeredge 212 of the airfoil 200 and a central hub of the turbine orcompressor.

The body 202 of the airfoil 200 can include a leading edge 208 at thefront of the airfoil 200 and a trailing edge 210 at the rear of theairfoil 200. A chord line is represented by a straight line connectingthe leading edge 208 and trailing edge 210 together. The chord length,c, is the length of the chord line.

The body 202 of the airfoil 200 can extend between the central hub,e.g., of a turbine, and the outer edge 212 of the airfoil 200. Whendisposed within the previously described turbofan jet engine 10, theouter edge 212 of the airfoil 200 can be spaced apart from a shroud ofthe turbofan jet engine 10 by a small clearance, as describedhereinafter. A plenum 214 can extend from the outer edge 212 of theairfoil 200 into the body 202 to create a void at the outer edge 212 ofthe airfoil 200. The void can define a volume, bounded by a planedefined by the outer edge 212. In the illustrated embodiment, the plenum214 is surrounded on all lateral sides by a sidewall defined by the body202.

The airfoil 200 is typically formed as a one-piece casting of a suitablesuperalloy, such as a nickel-, cobalt-, or iron-based superalloy, whichhas acceptable strength at the elevated temperatures of operation in agas turbine engine. The plenum 214 can serve to eliminate material innon-essential areas of the airfoil 200. This can result in reduction ofthe weight of the airfoil 200 which can have important benefits,including reductions both in operating stresses and material costs.Additionally, the plenum 214 may create one or more desirableaerodynamic characteristics with respect to the shroud of the gasturbine engine, as described hereinafter.

FIG. 3 is a cross-sectional view of the airfoil 200 as seen along LineA-A in FIG. 2 . As depicted in FIG. 3 , the airfoil 200 can have aradial direction 216 which defines a center line of the body 202 asmeasured normal to the central hub at a location from which the airfoil200 extends from. The radial direction 216 may be oriented generallyperpendicular, i.e., normal, to an inner surface of the shroud of thegas turbine engine when the airfoil 200 is disposed therein. In theillustrated embodiment, the radial direction 216 intersects the plenum214. The radial direction 216 is depicted as being normal to the outeredge 212 of the airfoil 200. In certain instances, the radial direction216 may be defined as a line perpendicular relative to the outer edge212 of the airfoil 200.

The plenum 214 can define a void disposed between a suction-sidesidewall 218 and a pressure-side sidewall 220 of the body 202. Thesuction-side sidewall 218 can correspond with the side of the plenum 214disposed nearest to the suction surface 204 while the pressure-sidesidewall 220 can correspond with the side of the plenum 214 disposednearest to the pressure surface 206. The suction-side sidewall 218 andpressure-side sidewall 220 can join together at opposite ends to formthe laterally bounded plenum 214.

In the depicted embodiment, the suction-side sidewall 218 includes twosegments—a first sidewall portion 222 and a second sidewall portion 224.The second sidewall portion 224 can be disposed adjacent to and radiallyoutside of the first sidewall portion 222. In an embodiment, the secondsidewall portion 224 can extend between the first sidewall portion 222and the outer edge 212.

In an embodiment, the first sidewall portion 222 and second sidewallportion 224 can be immediately adjacent to one another and meet at ajunction line extending into the page. In certain instances, thejunction line can be spaced apart from the outer edge 212 of the airfoil200 by a fixed distance, as measured along the length of the junctionline. In other instances, the distance between the junction line and theouter edge 212 of the airfoil 200 can vary along the junction line. Forexample, the first sidewall portion 222 can have a first radial height,as measured parallel with the radial direction 216, at a first locationof the plenum 214, and a second radial height, as measured parallel withthe radial direction 216, at a second location of the plenum 214,wherein the first and second radial heights are different from oneanother. To create the variable height of the first sidewall portion222, the height of the second sidewall portion 224 can inversely change,the depth of the plenum 214 can change, or both.

In an embodiment, the first and second sidewall portions 222 and 224 canbe angularly offset from one another by an angle, β. That is, the firstand second sidewall portions 222 and 224 can lie along lines, as viewedin cross section (e.g., as depicted in FIG. 3 ), that intersect oneanother at the junction line. By way of example, the first and secondsidewall portions 222 and 224 can be angularly offset from one anotherby an angle, β, of at least 1°, such as at least 2°, such as at least3°, such as at least 4°, such as at least 5°, such as at least 10°, suchas at least 15°, such as by at least 20°. In an embodiment, the secondsidewall portion 224 can be parallel, or substantially parallel, withthe radial direction 216 of the airfoil 200 while the first sidewallportion 222 can be angularly offset from the radial direction 216 by theaforementioned angular offset between the first and second sidewallportions 222 and 224. In an embodiment, the junction between the firstand second sidewall portions 222 and 224 can occur at a sharp interface.In another embodiment, the junction between the first and secondsidewall portions 222 and 224 can occur over a gradual transition, e.g.,a rounded corner as viewed in cross section.

The pressure-side sidewall 220 is illustrated in FIG. 3 with a generallylinear cross-sectional shape. The pressure-side sidewall 220 can be liealong a line parallel, or generally parallel, with the radial direction216. Alternatively, the pressure-side sidewall can taper toward or awayfrom the radial direction 216 at the outer edge 212. In anotherembodiment, the pressure-side sidewall 220 can have a non-linearcross-sectional shape, such as, for example, an arcuate cross-sectionalshape, a segmented cross-sectional shape (e.g., similar to thesuction-side sidewall 218), another non-linear cross-sectional shape, orany combination thereof. In the illustrated embodiment, a thickness ofbody 202, as measured between the pressure side 206 and thepressure-side sidewall 220, is shown as being generally constant. Inother embodiments, the thickness of the body 202, as measured betweenthe pressure side 206 of the airfoil 200 and the pressure-side sidewall220 of the plenum 214 can vary as measured along the radial height ofthe body 202. In an embodiment, the pressure-side sidewall 220, or aportion thereof, can be parallel, or generally parallel, as viewed incross section, with the second sidewall portion 224 of the suction-sidesidewall 218.

The airfoil 200 can include a flared portion 226 extending from thesuction side 204 of the airfoil 200. The flared portion 226 can extendfrom the body 202 of the airfoil in the axial direction (e.g., axialdirection A as shown in FIG. 1 ). In an embodiment, the flared portion226 can be integrally formed with the body 202 of the airfoil 200. Thatis, the body 202 of the airfoil 200 and flared portion 226 can bemonolithically formed from a single piece. A dashed line 228 is shown inFIG. 3 to represent where traditional airfoils ends along the suctionsurface 204 in the axial direction. The flared portion 226 can introduceadditional mass to the airfoil 200 along (i.e., adjacent to) the suctionsurface 204. In particular, the flared portion 226 can increase the massof the airfoil 200 at the furthest radially outer portion of the body202 (i.e., adjacent to the outer edge 212), thus having the greatesteffect on the moment of inertia of the airfoil 200.

In an embodiment, the flared portion 226 can extend along less than 30%of a radial length of the airfoil 200, as measured between the outeredge 212 and the central hub of the airfoil 200, such as less than 25%of the radial length of the airfoil 200, such as less than 20% of theradial length of the airfoil 200, such as less than 19% of the radiallength of the airfoil 200, such as less than 18% of the radial length ofthe airfoil 200, such as less than 17% of the radial length of theairfoil 200, such as less than 16% of the radial length of the airfoil200, such as less than 15% of the radial length of the airfoil 200, suchas less than 14% of the radial length of the airfoil 200, such as lessthan 13% of the radial length of the airfoil 200, such as less than 12%of the radial length of the airfoil 200, such as less than 11% of theradial length of the airfoil 200, such as less than 10% of the radiallength of the airfoil 200. In a more particular embodiment, the flaredportion 226 can extend along less than 8% of the radial length of theairfoil 200, such as along less than 6% of the radial length of theairfoil 200, such as along less than 4% of the radial length of theairfoil 200, such as along less than 2% of the radial length of theairfoil 200. As used herein, radial length is intended to refer to thelength of the airfoil 200 in the radial direction as would be seen withthe airfoil 200 operationally disposed within a gas turbine engine(e.g., the turbofan jet engine 10 depicted in FIG. 1 ). That is,reference to the radial direction is made with respect to an orientationof the airfoil 200 as part of a turbine or compressor within the gasturbine engine. In an embodiment, the flared portion 226 can extend fromthe outer edge 212 of the airfoil 200 toward the central hub. In a moreparticular embodiment, the flared portion 226 can be part of, i.e.,define a portion of, the outer edge 212 of the airfoil 200.

In an embodiment, the flared portion 226 can extend along a distanceless than the chord length, c, of the airfoil 200. That is, for example,referring again to FIG. 2 , the length, L_(FP), of the flared portion226, as measured along the suction surface 204 can be less than thechord length, c, of the airfoil 200. By way of example, L_(FP) can beless than 0.99 c, such as less than 0.98 c, such as less than 0.97 c,such as less than 0.96 c, such as less than 0.95 c, such as less than0.9 c, such as less than 0.85 c, such as less than 0.8 c, such as lessthan 0.75 c. In an embodiment, L_(FP) can be less than 0.7 c, such asless than 0.65 c, such as less than 0.6 c. In another embodiment, L_(FP)can be at least 0.1 c, such as at least 0.15 c, such as at least 0.2 c,such as at least 0.3 c, such as at least 0.4 c. In yet anotherembodiment, L_(FP) can be within a range of 0.1 c and 0.99 c, such as ina range of 0.25 c and 0.75 c. Flared portions 226 having lengths,L_(FP), less than the chord length, c, can be referred to as miniflares. The flared portion 226 may be readily discernable on mini flaresas the portion of the suction surface 204 disposed between the leadingand trailing ends 208 and 210 and the mini flare will have a moretraditional, i.e., non-flared, cross-sectional profile. In a particularinstance, the flared portion 226 can be equally spaced apart from theleading and trailing ends 208 and 210 of the airfoil 200. In anotherinstance, the flared portion 226 can be disposed closer to one of theleading and trailing ends 208 or 210. For example, the flared portion226 can be skewed by up to 100% towards one of the leading and trailingends 208 and 210.

In some embodiments, the flared portion 226 can define a linearcross-sectional surface profile, as seen for example in FIG. 3 . Thatis, an exposed portion of the flared portion 226 can lie along agenerally straight line as viewed in cross section. In otherembodiments, the flared portion 226 can have an arcuate cross-sectionalsurface profile, such as a flared profile having an increasingcross-sectional surface profile angle, as measured, for example, withrespect to the radial direction 216. In an embodiment, a tip 227 of theflared portion 226 may be sharp, e.g., an angled corner formed betweenthe outer edge 212 of the airfoil 200 and the axial surface of theflared portion 226. In another embodiment, the tip 227 of the flaredportion 226 can have a rounded or segmented cross-sectional profile.

In an embodiment, the flared portion 226 can elongate the outer edge 212of the airfoil 200, as measured in the surface normal direction ascompared to a traditional airfoil, by at least 0.1 mm, such as at least0.5 mm, such as at least 1 mm, such as at least 2 mm, such as at least 5mm. In another embodiment, the flared portion 226 can elongate the outeredge 212 of the airfoil 200, as measured in the axial direction ascompared to a traditional airfoil, by at least 10 mm, such as at least15 mm, such as at least 20 mm, such as at least 25 mm, such as at least30 mm, such as at least 40 mm, such as at least 50 mm, such as at least60 mm, such as at least 70 mm, such as at least 80 mm, such as at least90 mm, such as at least 100 mm. In an embodiment, a thickness of body202, T_(BS), as measured between the suction surface 204 and thesuction-side sidewall 218 at the outer edge 212 of the airfoil 200, canbe greater than the thickness of the body 202, T_(BP), as measuredbetween the pressure side 206 and the pressure-side sidewall 220 at theouter edge 212 of the airfoil 200. For example, T_(BS) can be at least1.01 T_(BP), such as at least 1.02 T_(BP), such as at least 1.03 T_(BP),such as at least 1.04 T_(BP), such as at least 1.05 T_(BP), such as atleast 1.1 T_(BP), such as at least 1.2 T_(BP), such as at least 1.5T_(BP), such as at least 2 T_(BP), such as at least 2.5 T_(BP), such asat least 3 T_(BP). In certain instances, T_(BS) may be no greater than10 T_(BP), such as no greater than 5 T_(BP).

In an embodiment, the flared portion 226 can define a height, H_(F), asmeasured in a direction perpendicular to the chord, that is greater thana largest axial dimension of the flared portion 226, A_(F), as measuredat the outer edge 212 of the airfoil 200. By way of example, H_(F) canbe at least 1.01 A_(F), such as at least 1.05 A_(F), such as at least1.1 A_(F), such as at least 1.25 A_(F), such as at least 1.5 A_(F), suchas at least 1.75 A_(F), such as at least 2.0 A_(F), such as at least 3.0A_(F), such as at least 4.0 A_(F), such as at least 5.0 A_(F), such asat least 10.0 A_(F), such as at least 20.0 A_(F).

The largest axial dimension of the flared portion 226, A_(F), can beless than the thickness of the body 202, T_(BS), as measured between thesuction surface 204 and the suction-side sidewall 218. By way ofexample, A_(F) can be less than 0.9 T_(BS), such as less than 0.85T_(BS), such as less than 0.8 T_(BS), such as less than 0.75 T_(BS),such as less than 0.7 T_(BS), such as less than 0.65 T_(BS), such asless than 0.6 T_(BS), such as less than 0.55 T_(BS), such as less than0.5 T_(BS), such as less than 0.45 T_(BS), such as less than 0.4 T_(BS),such as less than 0.35 T_(BS), such as less than 0.3 T_(BS), such asless than 0.25 T_(BS), such as less than 0.2 T_(BS).

In an embodiment, the suction surface 204 at the flared portion 226 canlie along a line angularly offset from the radial direction 216 of theairfoil 200 by an angle, α. The first sidewall portion 222 can beangularly offset from the radial direction 216 by the same, orsubstantially same, angle, α. The suction surface 204 at the flaredportion 226 can be generally parallel with the first sidewall portion222 of the suction-side sidewall 218, as measured at a same radialdistance from the outer edge 212. That is, T_(BS) can be approximatelyequal at multiple, radially spaced apart planes intersecting the firstsidewall portion 222 of the suction-side sidewall 218. This can occurwhere α is equal to β. Meanwhile, in accordance with an embodiment,T_(BS) can be different at multiple, radially spaced apart planesintersecting the second sidewall portion 224 of the suction-sidesidewall 218. For example, T_(BS) can increase as the plane ofmeasurement gets closer to the outer edge 212 of the airfoil 200. Incertain instances, the rate of change of T_(BS) relative to the radialdistance relative to the outer edge 212 can determine the relativeamount of flaring occurring at the flared portion 226.

Still referring to FIG. 3 , in certain instances, the flared portion 226can be identified by a relative angular displacement between the suctionsurface 204 and the pressure surface 206, as measured at anapproximately same location in the radial R direction (FIG. 1 ). Thepressure surface 206 can be angularly offset from the radial direction216 by an angle, γ, that is less than the angle α of the flared portion226. For example, α can be at least 5 degrees less than γ, such as atleast 10 degrees less than γ, such as at least 15 degrees less than γ,such as at least 20 degrees less than γ, such as at least 25 degreesless than γ, such as at least 30 degrees less than γ, such as at least35 degrees less than γ, such as at least 40 degrees less than γ.

FIG. 4 illustrates a cross-sectional view of the airfoil 200 as seen inaccordance with an embodiment. A shroud S of the gas turbine engine isshown to depict a relative spacing between the outer edge 212 of theairfoil 200 and the shroud S when the airfoil 200 is installed in a gasturbine engine. In the illustrated embodiment, the outer edge 212 of theairfoil 200 is spaced apart from the shroud S by a clearance, CLR. In anembodiment, clearance, CLR, may be measured at operating temperaturesand conditions, e.g., when the airfoil 200 is subjected to centripetalforces acting upon the rotating airfoil 200 which might cause the body202 of the airfoil 200 to elongate in the radial direction. In anembodiment, the clearance, CLR, can be at least 0.01 mm, such as atleast 0.05 mm, such as at least 0.1 mm, such as at least 0.5 mm, such asat least 1 mm, such as at least 1.5 mm, such as at least 2 mm, such asat least 3 mm, such as at least 4 mm, such as at least 5 mm. The height,H, of the plenum 214 is shown as measured from the outer edge 212 of theairfoil 200 to a radially inner surface of the plenum 214. The height,h, of the second portion 224 is shown as measured from the outer edge212. The height of the first sidewall portion 222 may be equal to theheight, H, of the plenum 214 minus the height, h, of the second portion224. In an embodiment, h/CLR is greater than 1.5, such as greater than2, such as greater than 2.5, such as greater than 3, such as greaterthan 3.5, such as greater than 4, such as greater than 4.5, such asgreater than 5. In this regard, the height of the second sidewallportion 224 can be greater than the clearance, CLR, between the outeredge 212 of the airfoil 200 and the shroud S. In certain instances,h/CLR can be impactful in affecting one or more aerodynamiccharacteristics of the airfoil 200, as described in greater detailhereinafter. While H/CLR may be impactful in affecting the aerodynamiccharacteristics of the airfoil 200, h/CLR being greater than 1.5 mayimprove the one or more aerodynamic characteristics while inclusion ofthe first sidewall portion 222 can minimize the resulting weight of theairfoil 200.

In an embodiment, the clearance, CLR, between the outer edge 212 of theairfoil 200 and the shroud S can be approximately uniform at alllocations along the outer edge 212 of the airfoil 200. In anotherembodiment, the clearance, CLR, can vary along the outer edge 212. Forinstance, the clearance of the outer edge 212 as measured at thepressure surface 206 can be different from the clearance of the outeredge 212 as measured at the suction surface 204 or another locationtherebetween.

FIG. 7 illustrates an enlarged view of an interface formed between theouter edge 212 of the airfoil 200 and an inner surface of the shroud Sas seen in circle B in FIG. 4 . Lines A₁, A₂, A₃, A₄, A₅, and A₆illustrate exemplary flow paths of air from the plenum 214 as the airpaths pass through the clearance, CLR, between the airfoil 200 andshroud S. The sharp edge 700 formed between the second sidewall portion224 of the suction-side surface 218 of the plenum 214 and the outer edge212 of the airfoil 200 can cause air passing between the airfoil 200 andshroud S to compress in the radial direction, generating the formationof an air gap G and a vena contracta effect where the air streamcross-sectional area is the smallest and the air velocity is thelargest. Generation of the air gap G at the sharp edge 700, therebycreating the vena contracta effect, can enhance the aerodynamiccharacteristics of the airfoil 200, improving performance andaerodynamic efficiencies of the gas turbine engine. In this regard,having the second sidewall portion 224 generally normal with the shroudS and/or radial direction 216 can be beneficial for engine performance.

In contrast, FIG. 5 illustrates an airfoil 500 without a plenum sidewallhaving an angle normal to the shroud S, as seen adjacent to the outeredge 212 of the airfoil 500. Instead, the plenum sidewall illustrated inFIG. 5 is generally linear and angularly offset from the radialdirection 216 as measured at the outer edge 212. FIG. 6 illustrates anenlarged view of the interface formed between the airfoil 500 and shroudS as seen in Circle C in FIG. 5 . As illustrated, the vena contractaeffect of the airfoil 500 is diminished as compared to that exhibited bythe airfoil 200 depicted in FIG. 7 as a result of the angularly offsetplenum sidewall. Notably, the air gap G is reduced in size, resulting ina larger cross-sectional area for air to pass through the clearancebetween the outer edge 212 of the airfoil 500 and the shroud S. In turn,engine performance is reduced as the airfoil 500 is less aerodynamicallyefficient.

Accordingly, use of a sharp edge 700 (FIG. 7 ) is advantageous increating a desirable aerodynamic characteristic. Referring again to FIG.4 , dashed line X depicts an exemplary view of the suction-side surface218 with the entire surface oriented parallel with the radial direction216. That is, unlike the suction-side surface 218 previously describedincluding the first sidewall portion 222 and second sidewall portion224, the dashed line X represents an example of the suction-side surface218 if the entire radial height H of the suction-side surface 218 of theplenum 214 were oriented parallel with the radial direction 216 whileaffording sufficient strength to the flared portion 226. When usingflared portions 226, the inclusion of such a linear side surface 218adds unnecessary weight to the airfoil 200 at the radially outermostportion there along, most affecting the centripetal forces acting on theairfoil 200. Specifically, the wall thickness between the suctionsurface 204 and the side surface 218 of the plenum 214 becomesunnecessarily large as a result of including the flared portion 226.

Accordingly, use of a dual-angled suction-side surface 218, as depictedin FIG. 4 and as previously described, including first and secondsidewall portions 222 and 224 can enhance aerodynamic performance whilereducing unnecessary weight associated with an entirely linearsuction-side surface 218 oriented parallel with the radial direction216.

FIG. 8 illustrates a view of an airfoil 800 in accordance with anembodiment herein as seen along the radial direction of the airfoil 800looking perpendicular to the outer edge 812 thereof. The airfoil 800 canhave any one or more features similar to the airfoil 200 previouslydescribed. The plenum 814, for example, is depicted extending betweenthe leading and trailing ends 808 and 810. Located adjacent to theplenum 814 is a recess 830 which extends from the outer edge 812 intothe airfoil 800. The recess 830 is shown disposed between the plenum 814and the flared portion 826. In an embodiment, the recess 830 extends adistance along the airfoil 200 less than a distance of the plenum 814.In another embodiment, the recess 830 and plenum 814 can have the samelengths, as measured between the leading and trailing ends 808 and 810.In yet another embodiment, the recess 830 can have a length greater thanthe length of the plenum 814.

FIG. 9 illustrates a cross-sectional view of the airfoil 800 shown inFIG. 8 as seen along Line B-B. The plenum 814 defines a depth D_(P), asmeasured from the outer edge 812, greater than a depth, D_(R), of therecess 830, as measured from the outer edge 812. In an embodiment, D_(P)can be at least 1.01 D_(R), such as at least 1.05 D_(R), such as atleast 1.1 D_(R), such as at least 1.25 D_(R), such as at least 1.5D_(R), such as at least 1.75 D_(R), such as at least 2.0 D_(R). Inanother embodiment, D_(P) can be no greater than 200 D_(R), such as nogreater than 100 D_(R), such as no greater than 25 D_(R), such as nogreater than 10 D_(R). In some instances, the depth, D_(R), of therecess 830 can be uniform, or generally uniform, along the entire recess830. In other instances, the depth, D_(R), of the recess 830 can varyalong the recess 830. For example, the recess 830 can have a firstdepth, D_(R1), at a first location, and a second depth, D_(R2), at asecond location different from the first location.

The recess 830 includes first and second sidewalls 832 and 834. Thefirst sidewall 832 is shown with an angularly offset surface as comparedto the radial direction of the airfoil 800 while the second sidewall 834is shown with a surface generally parallel with the radial direction ofthe airfoil 800. In another embodiment, the first and second sidewalls832 and 834 of the recess 830 can be parallel, or approximatelyparallel, with one another. In another embodiment, the second sidewall834 (closer to the plenum 814) can be angularly offset from the radialdirection of the airfoil 800 and the first sidewall 832 (closer to theflared portion 826) can be parallel with the radial direction of theairfoil 800.

In the illustrated embodiment, the first and second sidewalls 832 and834 lie along generally straight lines, as seen in cross section. Inanother embodiment, at least one of the first and second sidewalls 832and 834 can have an arcuate and/or multi-segmented sidewall contour. Inan embodiment, an angle formed between the first and second sidewalls832 and 834 can be at least 1°, such as at least 5°, such as at least10°, such as at least 15°. In another embodiment, the angle between thefirst and second sidewalls 832 and 834 can be no greater than 89°, suchas no greater than 60°, such as no greater than 45°.

The recess 830 can define an aspect ratio [L_(R):W_(R)] as measured by arelative length, L_(R), of the recess 830 as compared to a relativewidth, W_(R), thereof. In an embodiment, the aspect ratio can be betweenapproximately 500:1 and 2:1, such as between 100:1 and 2.5:1, such asbetween 50:1 and 3:1. The recess 830 can further define another aspectratio [L_(R):D_(R)] as measured by the relative length, L_(R), of therecess 830 as compared to a relative depth, D_(R), thereof. In anembodiment, the aspect ratio can be between approximately 100:1 and 5:1,such as between 750:1 and 10:1, such as between 500:1 and 20:1. Incertain instances, the recess 830 can enhance performance of the airfoil800, such as improve aerodynamic performance, along the outer edge 812.

FIGS. 10 and 11 illustrate an airfoil 1000 having any one or moresimilar characteristics as described with respect to the airfoils 200and 800 previously described herein. For example, the airfoil 1000 caninclude a plenum 1014 extending from a outer edge 1012 of the airfoil1000 and a flared portion 1026 at the suction surface 1004. FIG. 10 isan axial view of the airfoil 1000 as seen along the radial direction.FIG. 11 is a cross-sectional view of the airfoil 1000 as seen along LineC-C in FIG. 10 .

In certain embodiments, the airfoil 1000 can include a cooling cavity1036 disposed within the airfoil 1000. Use of a cooling cavity 1036 maybe particularly useful for stage one turbine blades which operate withinareas of the gas turbine engine having very high temperatures, e.g.,areas of the gas turbine engine located adjacent to the combustion area.In an embodiment, the cooling cavity 1036 can include a singlevolumetric cavity. In another embodiment, the cooling cavity 1036 caninclude a plurality of cooling cavities, such as a plurality ofinterconnected cooling cavities. The cooling cavity 1036 can be in fluidcommunication with the outer edge 1012 of the airfoil 1000 through oneor more cooling holes 1038. The one or more cooling holes 1038 canextend, for example, from the cooling cavity 1036 to the outer edge 1012and pass between the plenum 1014 and flared portion 1026. The one ormore cooling holes 1036 can define a length, L_(L), as measured betweenthe cooling cavity 1036 and the outer edge 1012. In an embodiment, allof the one or more cooling holes 1036 can have the same lengths, L_(L),as compared to one another. In another embodiment, at least two of theone or more cooling holes 1036 can have different lengths, L_(L), ascompared to one another. The one or more cooling holes 1036 can definecross-sectional sizes, e.g., diameters, and/or shapes as measured normalto the length, L_(L). In an embodiment, the one or more cooling holes1036 can all have the same cross-sectional sizes and/or shapes ascompared to one another. In another embodiment, at least two of the oneor more cooling holes 1036 can have different cross-sectional sizesand/or shapes as compared to one another.

In an embodiment, the one or more cooling holes 1038 can include asingle cooling hole 1036. In another embodiment, the one or more coolingholes 1038 can include at least two cooling holes, such as at leastthree cooling holes, such as at least four cooling holes, such as atleast five cooling holes, such as at least six cooling holes, such as atleast seven cooling holes, such as at least eight cooling holes, such asat least nine cooling holes, such as at least ten cooling holes, such asat least twenty cooling holes, such as at least fifty cooling holes. Inan exemplary embodiment with a plurality of cooling holes 1038, at leastthree of the cooling holes 1038, such as all of the cooling holes 1038,can be equidistantly spaced apart from one another. In another exemplaryembodiment with a plurality of cooling holes 1038, at least three of thecooling holes 1038, such as all of the cooling holes 1038 can be spacedapart from one another by different distances. For example, a firstadjacent pair of cooling holes 1038 can be spaced apart by a firstdistance different than a distance between a second adjacent pair ofcooling holes 1038.

As depicted in FIG. 11 , in accordance with one or more embodiments, atleast one of the one or more cooling holes 1038 can lie along lines thatare angularly offset from the radial direction of the airfoil 1000. Inother embodiments, at least one of the one or more cooling holes 1038can be parallel, or substantially parallel, with the radial direction ofthe airfoil 1000. In another embodiment, at least one of the one or morecooling holes 1038 can include a plurality of segmented portions, i.e.,a plurality of linear or arcuate sections joined together.

The one or more cooling holes 1038 can be in fluid communication withthe cooling cavity 1036 and provide cooled fluid, e.g., cooled air, fromthe cooling cavity 1036 to the outer edge 1012 of the airfoil 1000.

FIGS. 12 and 13 illustrate another embodiment of the airfoil 1000.Similar to the embodiment illustrated in FIGS. 10 and 11 , the airfoil1000 of FIGS. 12 and 13 include one or more cooling holes 1038. However,unlike the embodiment illustrated in FIGS. 10 and 11 , the cooling holes1038 terminate at the outer edge 1012 at an enlarged area 1040. In anembodiment, the enlarged area 1040 can include a cooling hole extensionsubstantially similar to the cooling hole 1038 but having a larger sizeand/or shape as compared to the underlying cooling hole 1038. By way ofexample, at least one of the one or more cooling holes 1038 can define afirst dimension, e.g., a first diameter, as measured at a first locationalong the length, L_(L), of the cooling hole 1038, and a seconddimension, e.g., a second diameter, as measured at a second locationalong the length, L_(L), of the cooling hole 1038 adjacent to the outeredge 1012, where the second size is greater than the first size. Forinstance, the second size can be at least 101% the first size, such asat least 105% the first size, such as at least 110% the first size, suchas at least 125% the first size. In another embodiment, the second sizeis no greater than 1000% the first size.

In the depicted embodiment, the enlarged area 1040 comprises a channelextending between the one or more cooling holes 1038. That is, theenlarged area 1040 extends continuously between adjacent cooling holes1038. In such a manner, the cooling holes 1038 can be in fluidcommunication with one another through the channel adjacent to the outeredge 1012 of the airfoil 1000. This can result in increased cooling andcirculation along the outer edge 1012 of the airfoil 1000. In anembodiment, the channel connecting the cooling holes 1038 can have aconsistent shape and/or size as measured along the channel. In anotherembodiment, at least one of the shape and size of the channel canfluctuate or vary at different locations along the channel. For example,the channel may have a minimum depth, as measured from the outer edge1012, at a midpoint between adjacent cooling holes 1038.

Airfoils are generally permitted to deform by a prescribed toleranceduring use. Such deformation can be created by materials operating athigh velocities, under high loads, at high temperatures. It is notuncommon for the outer edges of airfoils to occasionally come intocontact with the shroud. When such contact occurs, the use of coolingholes 1038 may be mitigated by scraped material from the outer edge ofthe airfoil closing the entrance to the cooling hole 1038. Use of anenlarged area 1040 at the outer edge 1012 of the cooling hole 1038 canprevent the cooling hole 1038 from closing shut if the airfoil 1000radially expands during operation and contacts the shroud S. That is,the enlarged area 1040 of the cooling hole 1038 at the outer edge 1012can allow for some degree of contact between the airfoil 1000 and shroudS without substantially, or entirely, closing the cooling hole 1038.

FIGS. 14 and 15 depict another embodiment of the airfoil 1000 similar tothose depicted in FIGS. 10 to 13 . The enlarged area 1040 is depictedwith a tapered profile, e.g., tapering from a size of the cooling hole1038 to an enlarged size, as measured at the outer edge 1012. Thetapered profile is shown having a curved taper. FIGS. 16 and 17illustrate an embodiment of the airfoil 1000 where the enlarged area1040 has a linear taper, as measured from the cooling hole 1038 to theouter edge 1012 of the airfoil 1000. Moreover, the embodimentillustrated in FIGS. 16 and 17 does not interconnect the enlarged area1040 of adjacent cooling holes 1038. Instead, each enlarged area 1040depicted in FIGS. 16 and 17 is part of a discrete cooling hole 1038. Itshould be understood that aspects of the exemplary embodimentsillustrated in FIGS. 10 to 17 can be combined in any number ofcombinations and are not limited to those combinations as explicitlyillustrated in the Figures.

In general, airfoils described herein can exhibit increased aerodynamicperformance without adding unnecessary weight to the outer edge thereof.In such a manner, the airfoil described herein can generate increasedperformance without decreasing engine efficiency or durability.

This written description uses examples to disclose the disclosure,including the best mode, and also to enable any person skilled in theart to practice the disclosure, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they include structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal languages of the claims.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

Embodiment 1. An airfoil for a gas turbine engine defining a radialdirection and an axial direction, the airfoil comprising: a flaredportion extending from a suction surface of the airfoil at an outer edgeof the airfoil along the radial direction; and a plenum disposed at theouter edge of the airfoil, the plenum having a suction-side sidewall anda pressure-side sidewall, wherein the suction-side sidewall has a firstsidewall portion adjacent to a second sidewall portion disposed outsideof the first sidewall portion along the radial direction, wherein thefirst sidewall portion defines a first angle with a radial direction ofthe airfoil as measured by an average angular offset between the radialdirection and the first sidewall portion, wherein the second sidewallportion defines a second angle with the radial direction of the airfoilas measured by an average angular offset between the radial directionand the second sidewall portion, and wherein the first angle is greaterthan the second angle.

Embodiment 2. The airfoil of any one of the embodiments, wherein theouter edge of the airfoil is configured to be spaced apart from a shroudof the gas turbine engine by a clearance distance, CLR, wherein thesecond sidewall portion has a radial height, h, and wherein h/CLR isgreater than 2.

Embodiment 3. The airfoil of any one of the embodiments, wherein asuction surface of the flared portion is angularly offset from theradial direction of the airfoil by a third angle as measured by anaverage angular offset between the radial direction and the surface ofthe flared portion, and wherein the third angle is approximately equalto the first angle.

Embodiment 4. The airfoil of any one of the embodiments, wherein theairfoil further comprises a cooling cavity disposed within a body of theairfoil, and wherein the cooling cavity is in fluid communication withthe outer edge of the airfoil through one or more cooling holesextending between the cooling cavity and the outer edge of the airfoil.

Embodiment 5. The airfoil of any one of the embodiments, wherein atleast one of the one or more cooling holes comprises a first size, asmeasured at a first radial location along the one or more cooling holes,and a second size, as measured at a second radial location along the oneor more cooling holes, the second radial location being radially outsidethe first radial location, and wherein the second size is greater thanthe first size.

Embodiment 6. The airfoil of any one of the embodiments, wherein thesecond radial location corresponds with the outer edge of the airfoil.

Embodiment 7. The airfoil of any one of the embodiments, wherein theflared portion extends less than 80% of a chord length, c, of theairfoil, as measured at the outer edge of the airfoil.

Embodiment 8. The airfoil of any one of the embodiments, wherein theflared portion is spaced apart from a trailing edge of the airfoil and aleading edge of the airfoil.

Embodiment 9. The airfoil of any one of the embodiments, wherein theflared portion has a tapered profile.

Embodiment 10. The airfoil of any one of the embodiments, wherein theouter edge of the airfoil defines a recess disposed between and spacedapart from the plenum and a suction surface of the flared portion, asmeasured at the outer edge of the airfoil.

Embodiment 11. A gas turbine engine defining a radial direction, the gasturbine engine comprising: a shroud; and a turbine positioned at leastpartially inward of the shroud along the radial direction, the turbinehaving a plurality of turbine rotor blades, wherein at least one of theplurality of turbine rotor blades comprises: a flared portion extendingfrom a suction surface of the airfoil at an outer edge of the airfoil;and a plenum disposed at the outer edge of the airfoil, the plenumhaving a suction-side sidewall and a pressure-side sidewall, wherein thesuction-side sidewall has a first sidewall portion adjacent to a secondsidewall portion disposed radially outside of the first sidewallportion, wherein the second sidewall portion is approximately normalrelative to an inner surface of the shroud, and wherein the firstsidewall portion is angularly offset from the second sidewall portion.

Embodiment 12. The gas turbine engine of any one of the embodiments,wherein the at least one of the plurality of turbine rotor bladescomprises all of the plurality of turbine rotor blades.

Embodiment 13. The gas turbine engine of any one of the embodiments,wherein the outer edge of the airfoil is spaced apart from the shroud bya clearance distance, CLR, wherein the second sidewall portion has aradial height, h, and wherein h/CLR is greater than 2.

Embodiment 14. The gas turbine engine of any one of the embodiments,wherein the first sidewall portion and second sidewall portion lie alongbest fit lines that intersect at a junction having a junction angle, β,of at least 5°.

Embodiment 15. The gas turbine engine of any one of the embodiments,wherein the flared portion extends no greater than 80% of a chord lengthof the airfoil, as measured at the outer edge of the airfoil, andwherein the flared portion is spaced apart from a trailing edge of theairfoil and a leading edge of the airfoil.

Embodiment 16. The gas turbine engine of any one of the embodiments,wherein the airfoil further comprises a cooling cavity disposed within abody of the airfoil, and wherein the cooling cavity is in fluidcommunication with the outer edge of the airfoil through one or morecooling holes extending between the cooling cavity and the outer edge ofthe airfoil.

Embodiment 17. An airfoil for a gas turbine engine defining a radialdirection and an axial direction, the airfoil comprising: a flaredportion extending from a suction surface of the airfoil along less than20% of a radial length of the airfoil, as measured from an outer edge ofthe airfoil along the radial direction, wherein the flared portionextends less than 80% of a chord length, c, of the airfoil, as measuredat the outer edge of the airfoil.

Embodiment 18. The airfoil of any one of the embodiments, wherein theairfoil further comprises: a body having a plenum disposed at the outeredge, the plenum having a suction-side sidewall and a pressure-sidesidewall, wherein the suction-side sidewall has a first sidewall portionadjacent to a second sidewall portion disposed radially outside of thefirst sidewall portion, wherein the second sidewall portion isapproximately parallel with a radial direction of the airfoil, andwherein the first sidewall portion is angularly offset from the radialdirection of the airfoil.

Embodiment 19. The airfoil of any one of the embodiments, wherein asuction surface of the airfoil is angularly offset from the radialdirection of the airfoil by an angle, α, as measured by an averageangular offset between the radial direction and the suction surface ofthe airfoil, and wherein the first sidewall portion is angularly offsetby an angle of approximately α, as measured by an average angular offsetbetween the radial direction and the first sidewall portion.

Embodiment 20. The airfoil of any one of the embodiments, wherein asurface of the flared portion is disposed at an angle, α, as measuredrelative to the radial direction, wherein a pressure surface of theairfoil is disposed at an angle, γ, as measured relative to the radialdirection at an approximately same relative height in the radialdirection as α, and wherein γ is less than α.

What is claimed is:
 1. An airfoil for a gas turbine engine, comprising:a body defining a pressure side and a suction side, the body extendingbetween a leading edge and a trailing edge and also extending along aradial direction to define an outer edge; a radially-extending plenumdisposed at the outer edge of the airfoil and having a suction-sidesidewall and a pressure-side sidewall; a flared portion disposed at theouter edge of the body and extending from the suction side; and a recessdisposed in the outer edge and positioned between the flared portion andthe suction-side sidewall of the plenum.
 2. The airfoil of claim 1,wherein the plenum defines a plenum depth with respect to the outeredge, and the recess defines a recess depth smaller than the plenumdepth with respect to the outer edge.
 3. The airfoil of claim 1, whereinthe recess comprises a first sidewall adjacent the flared portion and asecond sidewall adjacent the plenum.
 4. The airfoil of claim 3, whereinthe second sidewall extends radially.
 5. The airfoil of claim 3, whereinthe first sidewall is angularly offset from the second sidewall.
 6. Theairfoil of claim 3, wherein at least one of the first sidewall or thesecond sidewall has a curved geometric profile.
 7. The airfoil of claim3, wherein at least one of the first sidewall or the second sidewall hasa linear geometric profile.
 8. The airfoil of claim 1, furthercomprising a cooling cavity within the body and at least one coolinghole fluidly coupling the cooling cavity to the recess.
 9. The airfoilof claim 8, wherein the at least one cooling hole comprises multiplecooling holes each terminating at the recess, wherein the recess definesa channel fluidly coupling the multiple cooling holes.
 10. The airfoilof claim 9, wherein a cooling hole of the multiple cooling holescomprises a first size at a first radial location and a second size at asecond radial location, the second radial location being radiallyoutside the first radial location, and wherein the second size isgreater than the first size.
 11. The airfoil of claim 1, wherein theflared portion is spaced from each of the leading edge and the trailingedge.
 12. The airfoil of claim 1, wherein the flared portion extendsless than 80% of a chord length of the airfoil as measured at the outeredge.
 13. The airfoil of claim 1, wherein a surface of the flaredportion is disposed at an angle α and the pressure side is disposed atan angle γ, with respect to the radial direction, wherein the angle γ isless than the angle α.
 14. A gas turbine engine defining a radialdirection and comprising: a shroud; and a turbine positioned at leastpartially inward of the shroud along the radial direction, the turbinehaving a plurality of airfoils, with an airfoil in the plurality ofairfoils comprising: a body defining a pressure side and a suction side,the body extending between a leading edge and a trailing edge, and alsoextending along a radial direction to define an outer edge confrontingthe shroud; a radially-extending plenum disposed at the outer edge ofthe airfoil and having a suction-side sidewall and a pressure-sidesidewall; a flared portion disposed at the outer edge of the body andextending from the suction side; and a recess disposed in the outer edgeand positioned between the flared portion and the suction-side sidewallof the plenum.
 15. The gas turbine engine of claim 14, wherein eachairfoil in the plurality of airfoils comprises the flared portion, therecess, and the plenum.
 16. The gas turbine engine of claim 14, whereinthe plenum defines a plenum depth with respect to the outer edge, andthe recess defines a recess depth smaller than the plenum depth withrespect to the outer edge.
 17. The gas turbine engine of claim 14,wherein the recess comprises a first sidewall adjacent the flaredportion and a second sidewall adjacent the plenum, wherein at least oneof the first sidewall or the second sidewall has a curved geometricprofile.
 18. The gas turbine engine of claim 14, further comprising acooling cavity within the body and at least one cooling hole fluidlycoupling the cooling cavity to the recess.
 19. The gas turbine engine ofclaim 18, wherein the at least one cooling hole comprises multiplecooling holes each terminating at the recess, wherein the recess definesa channel fluidly coupling the multiple cooling holes.
 20. The gasturbine engine of claim 14, wherein the flared portion is spaced fromeach of the leading edge and the trailing edge.